A typical gas turbine engine includes in serial flow communication a compressor, combustor, high pressure turbine (HPT), and a power turbine or low pressure turbine (LPT). The HPT includes one or more rotor blade stages with each having a stationary turbine nozzle disposed upstream therefrom for channeling thereto combustion gas from the combustor. The HPT nozzle firstly receives the hot combustion gas from the combustor and channels it to the first stage rotor blades which extract energy therefrom. The pressure and temperature of the combustion gas are decreased when the gas exits the first stage blades and enters the second stage turbine nozzle from which it flows to the second stage blades which extract further energy therefrom.
Since the turbine nozzles channel the hot combustion gas, they are cooled using a portion of air bled from the compressor to prevent thermal distress thereof and for effecting a useful life during operation. They require maximum cooling during high power operation of the engine such as during takeoff for an aircraft gas turbine engine, and reduced cooling at low power operation such as during cruise. The air bled from the compressor which is used for nozzle cooling is therefore not used in the combustion process and correspondingly decreases the overall efficiency of the engine.
Accordingly, it is desired to use as little bleed air as possible for effectively cooling the nozzles for maximizing the operating efficiency of the engine. For example, it is conventionally known to modulate the bleed airflow to the second stage turbine nozzle for providing maximum airflow at high power operation and minimum airflow at low power operation. This may be accomplished by providing a conventional modulating valve in the bleed line between the compressor and the second stage nozzle. However, the minimum bleed air flowrate must nevertheless be suitably high and have a suitable pressure to prevent backflow of combustion gases from the main flowpath into the nozzle vane.
More specifically, the second stage nozzle includes a plurality of circumferentially spaced apart vanes typically formed in arcuate segments each having two vanes extending between outer and inner bands. The vanes are hollow and typically include an impingement air baffle therein for enhancing cooling of the vanes, with the bleed air being introduced through the radially outer band for flow radially inwardly through the vanes for discharge in part through the radially inner bands. A conventional thermal shield is typically disposed radially inwardly of the inner bands and extends axially between the first and second stage rotors. A conventional honeycomb seal is joined to the inner band and cooperates with a plurality of radially outwardly extending labyrinth seal teeth extending outwardly from the thermal shield. This arrangement defines a forward cavity or chamber between the first stage rotor and the second stage nozzle, and an aft cavity or chamber disposed between the second stage nozzle and the second stage rotor. The inner band includes forward and aft purge holes through which a portion of the bleed air is discharged into the respective forward and aft chambers from the vane. The purge air has a suitable pressure and flowrate for flowing from the respective forward and aft chambers radially outwardly into the main combustion flowpath to continually purge the chambers and prevent backflow of the combustion gases into these chambers.
Accordingly, the bleed air to the vanes must be suitably modulated to provide a minimum flowrate and pressure to effectively continually purge the forward and aft chambers and prevent backflow into these chambers or into the vanes themselves. However, this minimum flowrate typically required for obtaining suitable purge flow is also typically more than is required for effectively cooling the nozzle vanes during low power operation such as cruise. Overall efficiency of the engine is, therefore, being reduced to ensure effective purge flow without backflow.